Satellite System

ABSTRACT

A satellite system operates at altitudes between 100 and 350 km relying on vehicles including a self-sustaining ion engine to counteract atmospheric drag to maintain near-constant orbit dynamics. The system operates at altitudes that are substantially lower than traditional satellites, reducing size, weight and cost of the vehicles and their constituent subsystems such as optical imagers, radars, and radio links. The system can include a large number of lower cost, mass, and altitude vehicles, enabling revisit times substantially shorter than previous satellite systems. The vehicles spend their orbit at low altitude, high atmospheric density conditions that have heretofore been virtually impossible to consider for stable orbits. Short revisit times at low altitudes enable near-real time imaging at high resolution and low cost. At such altitudes, the system has no impact on space junk issues of traditional LEO orbits, and is self-cleaning in that space junk or disabled craft will de-orbit.

CROSS REFERENCE TO RELATED APPLICATIONS

This application claims priority to U.S. Provisional Patent ApplicationSer. No. 62/430,727, filed Dec. 6, 2016, entitled “A Satellite System”.The entirety of U.S. Provisional Patent Application Ser. No. 62/430,727is incorporated herein by reference.

BACKGROUND

Satellites are used in many aspects of modern life, including earthobservation and reconnaissance, telecommunications, navigation (e.g.,global positions systems, or “GPS”), environmental measurements andmonitoring and many other functions. A key advantage of satellites isthat they remain in orbit due to their high velocity that creates anoutward centripetal force equal to gravity's inward force. Therefore,once in orbit, they stay there typically for years or decades. See, forexample, FIG. 11, which graphically illustrates a best and worst casecurve for expected lifetime of orbiting vehicles as a function ofaltitude. Since the velocities are so high (e.g., 3-8 km/s, depending onaltitude), atmospheric drag should be minimized and/or avoided, whichmeans they typically operate outside virtually any trace of theparticles that constitute the atmosphere. In addition to drag,atmospheric collisions with particles, even at trace concentrations, canoverheat, damage or eventually destroy the satellite.

Satellites are therefore clearly differentiated from atmospheric flying(i.e., airborne) vehicles such as airplanes, unmanned aerial vehicles(UAVs), helicopters or balloons in which the atmosphere supports liftand the vehicles operate at velocities typically between zero (i.e.,hovering) to 1-3 times the speed of sound and at altitudes below about35 km.

Satellite orbital heights are typically categorized in three broadsegments: low earth orbit (LEO), medium earth orbit (MEO) andgeostationary earth orbit (GEO). The general uses and characteristics ofthese orbits are shown in Table I and represent generally accepted usageof the terms LEO, MEO and GEO. Satellites can orbit at any altitudeabove the atmosphere, and the gaps in altitude shown in Table 1, such asbetween LEO and MEO, are also used, if less regularly. It is also commonthat satellites may orbit in eccentric, non-circular orbits, therebypassing through a range of altitudes in a given orbit.

TABLE I Typical characteristics of common orbits. Altitude, Velocity,Orbit km km/s Exemplary Uses Comments LEO 400- 6.9-7.8 Earth Random2,000 observation, orbits, 3-10 sensing, ISS, Y lifetime, space telecomjunk issue, constellations little radiation MEO 15,000- 3.5 GPS, Highestradiation 20,000 GLONASS, (Van Allen Belt), Earth equatorial toobservation polar orbits GEO 42,000 3.1 Sat TV, high Can remain above BWtelecom, same spot on weather Earth, typically satellites equatorialorbits

For most satellites, their useful lifetime is determined by multiplefactors. For example, in the case of GEO satellites, small fluctuationsin solar winds and earth's gravity require regular use of fuel tomaintain the satellite's position and attitude. Once exhausted of fuel,a satellite is typically rendered useless and decommissioned. However,due to GEO height, such a satellite itself will stay in orbit virtuallyforever due to its altitude and near zero atmospheric drag. Due to theirapparent stationary position as viewed from earth's surface, they arewidely used for telecommunications and satellite TV. Their largedistance from Earth limits their usefulness in telephone services (timedelay) and in high-resolution imaging (distance). They encounter solarwinds and cosmic radiation that force use of very specialized andexpensive electronics to survive.

MEO satellites are in the mid-range, mostly similar to GEO satellitesexcept that they do not appear stationary when viewed from earth'ssurface. Their most common usage is for satellite positioning services,such as GPS, and certain Earth observation missions for which theirtrade-off in altitude between GEO and LEO is beneficial. Due to thepresence of the so-called Van Allen Belts, these satellites can sufferlarge amounts of radiation and therefore require very specialized andexpensive electronics to survive.

LEO satellites, conversely, may be in a constant state of very slightatmospheric drag requiring either regular boost to their altitude (e.g.fuel burns) or an end-of-useful-life caused by reentry and burn upsimilar to a meteor entering the earth's atmosphere. As an example, theInternational Space Station (ISS), orbiting at about 425 km, losesapproximately 2-4 km/month of altitude and requires regular fuel burnsto ensure it stays in proper orbit. But the atmospheric drag is stillvery low and LEO satellites can remain in orbit for years without fuelburns.

This relatively long life is the source of so-called “space junk”, inwhich any orbiting device can potentially collide with a usefulsatellite, thereby damaging or destroying it and creating additionalorbiting objects. It is a widely recognized issue that at some densityof space junk, probabilities of collisions increase, eventually leadingto a virtually unusable orbit. A beneficial element of the currentinvention is to provide satellite services without increasing the spacejunk issue and furthermore to enable a mechanism that will be“self-cleaning” in the chosen orbits of 100-350 km.

Due to various shielding effects, especially of earth's magnetic fields,LEO satellites encounter little radiation and therefore do notnecessarily require specialized and expensive electronics to survive. Anexception to this rule is the so-called South Atlantic Anomaly, or SAA,which is a region in which a higher density of energetic particles maybe found, causing short term interruptions of some electronics. Thiseffect can be mitigated by many known techniques, so does not present alarge issue for LEO satellites.

In fact, continual improvement in system operation is realized since bylowering the operating altitude, system components (e.g. optics,electronics, synthetic aperture radar (SAR), required solar panel area,etc.) can be made smaller, which in turn reduces vehicle size and drag,thereby enabling an even lower operating altitude, and so-on. While itis desirable to be closer to earth's surface (or any celestial body'ssurface, say Mars), atmospheric density effectively sets a lower limiton orbital altitude; or forces expensive, heavy counteracting systemssuch as on Gravity field and steady-state Ocean Circulation Explorersatellite (GOCE), discussed below. For bodies without an atmosphere,such as earth's moon, there is no lower limit other than hitting thebody itself.

SUMMARY

In one example, a satellite network is described that includes aplurality of satellites arranged in an orbit having an altitude between100 km and 350 km. Each satellite includes a system to ingest ambientair particles, to thermalize, concentrate and slow the incoming ambientair particles. Each satellite includes an ion engine configured toingest the ambient air particles, ionize the ambient air particles, andgenerate thrust from the ionized ambient air particles sufficient tomaintain the orbit of the satellite.

In another example, a satellite is configured to orbit a terrestrialsurface at an altitude between 100 km and 350 km. The satellite includesa system with an inlet to thermalize, concentrate and slow incomingambient air particles. The satellite also includes an ion engineconfigured to ingest and ionize the ambient air particles and generatethrust from the ionized ambient air particles sufficient to maintain theorbit of the satellite.

In yet another example, a method is described of generating thrust in aself-sustaining low earth orbit satellite. The method includes ingestingambient air particles by a system configured to slow the ambient airparticles by at least two orders of magnitude and to concentrate theambient air particles by at least one order of magnitude within saidsatellite. The method further includes ionizing the ambient airparticles by an ion engine, and accelerating the ionized ambient airparticles through an ejection port of said satellite to generate thrust.

BRIEF DESCRIPTION OF THE DRAWINGS

FIG. 1 shows an example satellite in accordance with aspects of thisdisclosure.

FIG. 2 shows an example satellite in accordance with aspects of thisdisclosure.

FIGS. 3 through 5 show perspective views of example satellites inaccordance with aspects of this disclosure.

FIG. 6 shows example concentrating and slowing system in accordance withaspects of this disclosure.

FIG. 7 shows another example concentrating and slowing system inaccordance with aspects of this disclosure.

FIG. 8 shows a cross-sectional schematic of an example concentrating andslowing system in accordance with aspects of this disclosure.

FIG. 9 shows a magnified view of an interaction with an exampleconcentrating and slowing system in accordance with aspects of thisdisclosure.

FIG. 10 shows an example of satellite necklaces in accordance withaspects of this disclosure.

FIG. 11 illustrates graphical data of traditional satellite operationalcapabilities.

FIG. 12 illustrates graphical data of density of air versus altitude.

FIG. 13 illustrates graphical data of density of air particles within anexample concentrating and slowing system in accordance with aspects ofthis disclosure.

FIG. 14 illustrates graphical data of gas properties within an exampleconcentrating and slowing system in accordance with aspects of thisdisclosure.

FIG. 15 is a flow chart of a method of generating thrust in aself-sustaining low earth orbit satellite, in accordance with aspects ofthis disclosure.

The several figures provided here describe examples in accordance withaspects of this disclosure. The figures are representative of examples,and are not exhaustive of the possible embodiments or full extent of thecapabilities of the concepts described herein. Where practicable and toenhance clarity, reference numerals are used in the several figures torepresent the same features.

DETAILED DESCRIPTION

This detailed embodiment is exemplary and not intended to restrict theinvention to the details of the description. A person of ordinary skillwill recognize that exemplary numerical values, shapes, altitudes,applications of any parameter or feature are used for the sole purposeof describing the invention and are not intended to be, nor should theybe interpreted to be, limiting or restrictive.

The current disclosure relates to vehicles operating at altitudesbetween about 100 km to 350 km, what is defined as a Near Earth Orbiter(NEO), using self-sustaining ion engines for orbiting where atmosphericdensity is too high for traditional satellites and too low for airbornevehicles. To be clear, a self-sustaining ion engine (SSIE) is an ionengine system that “scoops” in ambient atmospheric particles (e.g.,atoms and molecules), ionizes them, accelerates and then ejects them tocreate thrust on the vehicle. To remain in stable orbit, the SSIEgenerates sufficient thrust to overcome the vehicle's drag. The SSIE maybe powered by solar energy and require no stored propellant (as istypical for conventional ion engines) or other stored energy source,hence the name “self-sustaining.” Ion engines are also known as electricpropulsion (EP), or ion thrusters.

FIG. 1 illustrates an exemplary version of a NEO vehicle 100. The NEOvehicle 100 can include a concentrating and slowing (C&S) System, withan opening 102 to ingest and direct ambient air particles (see also,e.g., FIGS. 6-9). The NEO vehicle 100 can further include an ion engine104 to ionize and expel the air particles, thereby generating thrust tomaintain the desired orbit. One or more stabilization surfaces 106 and108 can be employed, designed to enhance the stability of the NEOvehicle 100, as well as support solar paneling to collect power. Theexample of FIG. 2, shown in perspective view with internal details ofthe NEO vehicle 100 revealed, illustrates the interior of an example C&Ssystem, where the opening 102 is directed to a narrowing channel 112 todirect ambient air particles into the ion engine 104. In the example ofFIG. 2, there are two openings 102 and two channels 112 feeding to acentral ion engine 104. In FIG. 2, the channels 112 are not directlyconnected to the ion engine 104, rather the channels 112 lead to exitports 110 at the back of the NEO vehicle 100. However, a connectionbetween the exit port 110 of the channel 112 is implied, and a person ofaverage skill will realize that a number of connection strategies couldbe used, depending on the number of ion engines, and the relativegeometry of the channels and ion engine(s). In other examples, a singleopening 102 and single channel 112 can be used, as well as three or moresuch openings and channels. Similarly, each channel 112 can feed into asingle, dedicated ion engine 104 or additional ion engines (not shown)may be fed by the channels 112. Moreover, in this example, the channel112 is oriented orthogonally from the bottom surface of the NEO vehicle100 (i.e., in the plane of surface 106). In other examples, the channel112 can be oriented parallel to the bottom surface of the NEO vehicle100, or any angle there between. In some examples, the channel 112 canhave a varying cross-sectional area (e.g., cylindrical, conical, etc.),or some alternative geometry, that effectively collects, slows anddirects the ambient air particles into the ion engine 104.

The NEO vehicle 100 of FIG. 2 further illustrates one or more radar orradio components 114 and one or more optical imagers 116, 118 (e.g.,variable field of view, multispectral imaging, etc.). Additional andalternative components may be included in the NEO vehicle 100, such asradio frequency (RF) antennae, sensors, electronics bays for electronicsand control circuitry, cooling, navigation, attitude control, and othercomponentry, depending on the conditions of the orbiting environment(e.g., air particle density), the particular application of thesatellite (e.g., optical imaging, telecommunications transceiver,scientific research etc.), for instance. The NEO vehicle 100 can furtherinclude energy storage capacity, such that the solar paneling canrecharge a battery to, for example, power the components (e.g., 114-118)and ion engine 104 of the NEO vehicle 100. FIGS. 3-5 illustrateadditional perspective views of the NEO vehicle 100.

As described below, a properly designed NEO vehicle enables an SSIE withsufficient thrust to ensure stable orbital operation by ingestingneutral air particles; concentrating and slowing them; then ionizing andaccelerating them, thus creating sufficient thrust to overcome thevehicle's drag. This enables Near Earth Orbiters, NEOs, a term we use todescribe the system and its constituent vehicles (i.e., a “NEO satellitesystem”, “NEO vehicle” or a “NEO satellite”) operating in stable orbitsat 100-350 km without carrying a store of primary propellant. Therefore,it is a purpose of this invention to describe a satellite system basedon orbital vehicles operating in stable Earth orbits at altitudes wellbelow traditional satellites, specifically between approximately 100 and350 km.

In the described examples, atoms and molecules are collected,concentrated, and ionized. Ionization occurs when at least one electronis stripped from an unionized atom or molecule, thereby creating an ionand free electron(s). The ions are then accelerated by the ion engine(104) to produce thrust, and the electrons are ejected from the NEOvehicle 100 to avoid charging effects on the satellite. It is an elementof the NEO vehicle 100 that it has an electron ejection mechanism toneutralize such charging.

Some example satellites orbit at altitudes below LEO (i.e., lower than350 km). Due to atmospheric drag at these altitudes, thrust is providedon a continuous or near-continuous basis or the vehicle's orbit willdecay in a matter of days, weeks or months, depending on altitude (see,e.g., FIG. 1). The NEO vehicle 100 described herein is configured toprovide sufficient thrust to maintain orbits between 100-350 km and todo so without having to carry propellant (e.g., ejected material thatcauses thrust) or fuel (e.g. stored energy source carried into orbitsuch as chemical fuels).

At altitudes lower than LEO, atmospheric density increasesexponentially, as shown in FIG. 12. Below an altitude of about 120 km,atmospheric density that causes atmospheric drag increases by an orderof magnitude about every 20 km. Meanwhile, above that breakpoint and upto about 400 km, the atmosphere changes by an order of magnitude aboutevery 50-75 km. The key effect is that atmospheric density, andtherefore drag, is about 5 orders of magnitude higher at an altitude of100 km compared to the altitude of the ISS at about 425 km. This is onereason maintaining a stable orbit for a traditional satellite withoutusing substantial fuel to create compensating thrust has not beenpossible. Accordingly, very few satellites operate below about 400 km,and those that do are often in highly elliptical orbits, thus spendingvery little time at the lower altitudes.

Maintaining a lower altitude orbit is desirable for multiple reasons.For instance, any earth imaging application can get effectively higherresolution images from a smaller, less complex camera simply by beingcloser to the surface. For example, if an imager is 3 times closer toits object, it will get approximately 9 times better resolution (i.e.,in pixels per area) for a given optical system. Similarly, fortelecommunications, due to the inverse square law relationship betweenradio frequency (RF) energy and distance, a transmitter that is 3 timescloser will create 9 times stronger signal at a receiver, or require 9times less power to achieve the same signal power at the receiver.Additionally, for an active radar application, being 3 times closerrequires 81 times less power for equivalent performance due to the 1/r⁴power law of radar. All of these factors enable the exemplary NEOvehicle 100 to reduce the size and cost of a NEO satellite systemsufficiently to enable large satellite constellations that have shortrevisit times at affordable cost.

In the example NEO vehicle 100, due to the immense savings in power, aNEO constellation employing radar applications may create near-real timeradar imagery of the earth's surface. Considering a synthetic apertureradar (SAR) as an example, typical satellite-based SAR systems in LEOorbits require average transmit powers in the kilowatt range. Suchradars therefore require very large solar arrays to power them and thencomplex cooling systems to remove the waste heat.

For a NEO SAR system with an 81 times reduction in power, the averagetransmit power consumption is reduced from, for example, 1 kW to about12 W. The solar panel size, weight and cooling required would alsoreduce by 81 times, thereby making such SAR systems that much cheaper tolaunch and operate. If the relative altitude were to be ¼ instead of ⅓of the traditional altitude, the savings would increase to 256 times andthe SAR example above may require less than 5 watts of transmittedpower.

The value and opportunity for this ultra-low power NEO SAR is that suchradars can image the earth's surface at night, through clouds, and eventhrough some dust storms. Therefore, a given NEO SAR system would beable to create useful images approximately 100% of the time whilereducing the statistical impact of night and cloud cover.

To achieve a SAR, an array of transmit/receive elements is provided withprecise spacing, typically at half the wavelength of the transmittedenergy. Such elements could be provided on a single NEO satellite withthe array attached to or trailing behind the NEO satellite in thedirection of motion, thereby creating the typically oblong beam patternrequired for SAR. The element array could also be created by a formationof NEO satellites which maintain accurate spacing, with such anarrangement also useful for longer wavelength radars. In both cases, therelative power savings is maintained due to the low altitude of the NEOorbits.

Other Earth observation requirements also benefit from lower altitudeorbits. For example, the European Space Agency GOCE satellite configuredto provide highly accurate gravitational measurements, was placed in avery low orbit (i.e., about 250 km) that normally would have decayedvery quickly. To stay in orbit for the desired 3-year life, thesatellite carried an ion thruster to expel its stored Xe atoms, therebycreating sufficient thrust to counteract the atmospheric drag. Launchingsufficient Xe into orbit was both expensive and heavy. And, when thesatellite ran out of Xe, initiating ion thrust was not possible, areentry process started that eventually destroyed the satellite.

Ambient Ingesting Ion Engines have also been proposed, also called agridded ion engine (GIE). Such an engine was never actuallydemonstrated, largely because the proposed approach has no fewer thantwo fatal flaws. For instance, a proposed ion engine relied on ingestingand accelerating ambient ions as opposed to neutral atoms. Previousattempts failed to realize that ingesting ions in equilibrium with therest of the atmospheric particles, and then accelerating them, wouldcreate no net thrust. Experiments have shown that such devices generateno net thrust due to this electrostatic equilibrium issues.

Further, the density of ions at these altitudes is about 3-5 orders ofmagnitude lower than neutral atoms or molecules. But neutral atoms andmolecules cause drag so the thrust generated exclusively from naturallyoccurring ions would inherently be many orders of magnitude less thanthe drag generated from neutral particles. Hence, the proposed GIE wouldnot have created sufficient thrust and therefore cannot be used for thecurrent, or any, NEO satellite application.

The example NEO satellite system described herein is capable ofproviding satellite imaging, communication services, radar imaging,earth measurements and other satellite services based on one or more NEOorbiting vehicles operating in long term, stable orbit at altitudesbetween approximately 100-350 km. Further, the satellite system includesan array of such NEO satellites in sufficient density to enablenear-real time coverage of the earth. Benefits of the NEO vehicle 100with a sustainable orbit would accrue to virtually all other satelliteapplications, such as communications.

The NEO vehicle 100 of FIG. 1 is capable of operating in a long term,stable, self-sustaining orbit at altitudes between approximately 100-350km without the need to carry either propellant or a stored energy sourceinto orbit. As described in greater detail below, the NEO vehicle 100 isable to ingest ambient (e.g. neutral) atmospheric atoms and molecules,ionize, accelerate and emit those atoms and molecules, and neutralizeany resulting charge built up on the NEO vehicle 100 in the process.

At altitudes of 100-350 km, earth's atmosphere is made up primarily ofO, O₂, N and N₂ in their neutral (i.e., un-ionized) state. In an examplewhere the NEO vehicle 100 orbits the Earth at about 160 km, the NEOvehicle 100 has an orbital velocity of about 7.8 km/sec, these atom andmolecule species have a velocity relative to the vehicle of the same 7.8km/sec. The overall atmospheric density (i.e., all species combined) isapproximately 1e-9 kg/m³ (i.e., an atmospheric pressure of about 1e-9atm). These conditions are not suitable for operating traditional ionengines.

To enable the ion engine 104 for use in a NEO vehicle 100, the relativevelocity of the particles should be slowed significantly and the densityincreased significantly, as is common with the operating pressures ofionization chambers in traditional ion engines. One element of the NEOsatellite system is the use of a system for concentrating and slowingthe atmospheric atoms and molecules. An exemplary concentrating andslowing system (C&S system) is shown and described in FIGS. 6-9.

FIG. 6 provides a cross-sectional view of the features of the C&SSystem. As shown in detail in FIG. 8, atmospheric particles enter thechamber opening 120, they encounter an inner wall 120 of the C&S system,which can be made of carbon, metal, ceramics, or other suitablematerials. On a microscopic scale, these materials have a surfaceroughness substantially larger than the size of the air particles. Theincoming particles are momentarily trapped by the surface roughness,thereby giving up their kinetic energy to the wall material, thencedeparting from contact with the surface having significantly lesskinetic energy. The particles are prone to reflection from the surfaceat random angles, resulting in so-called diffuse reflection, shown indetail in FIGS. 8 and 9.

The trapped particles do not attach to the surface 120, but are justdetained momentarily and thermalized (i.e., the particles “rattlearound” in the rough areas until they come to thermal equilibrium, orare “thermalized”, with the surface material). Once thermalized, theparticles are emitted from the wall 120 material having a new thermalvelocity, which is lower than their original incoming velocity and inrandom directions due to diffuse reflection. The emitted particles maythen hit other walls 122, with some fraction of the particles bouncingback out of the C&S opening 102, but many of them concentrating in thechannel 112 of the C&S system. At the opening 102, the air density is ata minimum, whereas within the channel 112, air density is several ordersof magnitude higher, as shown graphically in FIG. 13.

In one example, different wall materials are provided at differentlocations along the C&S system. For example, atomically smooth materialssuch as sapphire or other polished materials may be used to directincoming particles (i.e., at surface 122, or any other desirable surfacelocation), where interactions would be similar to specular reflection.The C&S system may then direct particles to a focal point within the C&Ssystem having a rougher surface (e.g., surface 126 of FIG. 8), therebythermalizing the particles. In such a structure, particles may befocused and then thermalized, increasing the percentage of particlescollected. Additionally or alternatively, catalytic materials may beused for the inner walls of the C&S system. Such materials couldefficiently recombine 0 atoms into O₂ molecules through surfacecatalytic reactions, in a process similar to the catalytic convertertechnology used in automobiles. Converting O into O₂ within the C&Ssystem could be beneficial, since it is desirable to ionize andaccelerate particles with a large molecular weight for maximum thrust.There are a number of material choices that could accomplish this. Theoverall geometry and choice of materials of the C&S system may beoptimized for various altitude and flight conditions.

FIGS. 6-8 show example configurations for the C&S system, includingopening 102 (i.e., a wide scoop) that feeds into channel 112 (i.e., anarrow duct). In one example shown in FIG. 7, the opening 102 containsone or more flat fins 124 that may act as a trap to prevent particlesfrom escaping the C&S system. FIG. 8 shows the effects of multipleinteractions between atoms and molecules and one or more surfaces of theC&S system. As shown in FIGS. 8 and 9, incoming particles 132 can“bounce around” off of walls 120, 122, 126, which each can be configuredwith a different material to optimize the trajectory of particles. Forexample, a fraction of incoming particles 132 may not initially enterthe duct, rather they may bounce back in the upstream direction andpossibly escape the mouth (i.e. opening 102) of the C&S system. Theaddition of fins 124 may prevent such particles from escaping andreflect a portion of them back into the duct. As shown in FIGS. 8 and 9,due to their random direction, escaping particles will likely collidewith a fin 124, these particles may undergo diffuse reflection, causingthe deflected particle 134 to eject at random angles 136 from thesurface of the fin 124, as shown in FIG. 9. Therefore a fraction of theescaping particles may escape, however, a fraction will be reflectedback towards the duct, and after further diffuse collisions, suchparticles may enter the duct as shown in FIG. 8. As shown in FIG. 9, theincoming particles 132 undergo diffuse reflection after contact with thesurface of fin 124, meaning that they bounce from the surface at randomangles 136 relative to their incoming track, which differs from, forexample, specular reflection experienced by light beams reflecting froma mirror.

Under favorable conditions, the trap can achieve an increase in densityand pressure in the C&S system compared to the duct and scoop geometrieswithout fins 124. However, if the angle of attack of the vehicle isnon-zero, or erosion of the thin fins 124 occurs due to high velocitycollisions with massive particles, the fins 124 may serve to block theincoming gas. Selection of material, size, and spacing of the fins 124can be customized for specific atmospheric conditions at a desiredaltitude for deployment of the NEO vehicle.

Various elements of the C&S system may be optimized, such as the precisegeometry of the opening 120 and channel 112, specular and diffusereflecting materials and their respective placement, as well as thegeometry of the fins 124, depending in accordance with expectedatmospheric and orbital conditions.

Optimization may include a throttling capability, such as a valveconfiguration at various points within the C&S system to build up gasduring high density portions of the orbit and use this excess gas inlower density portions of the orbit. A useful aspect of the NEO vehiclesis that even a simple inlet geometry, designed for purely diffusereflection, is able to produce gas conditions within the operating rangeof existing ion engines.

Referring to FIGS. 13 and 14, results are shown from direct simulationMonte Carlo (DSMC) calculations under actual atmospheric conditions at160 km of an example C&S system. DSMC modeling is useful, not only dueto the complex geometry of the C&S system, but because large densityincreases may result in gas conditions where particle-particlecollisions become significant and the flow is no longer free-molecularwithin portions of the C&S system. For instance, the C&S system canincrease atmospheric density and therefore pressure by over 200 timesthe atmospheric pressure at the entrance of the opening. The pressure atthe entrance of the duct 112 has increased to approximately 0.025 Pa (25mPa), which is a pressure that allows the ion engine 104 to operate.Some example engines have an operating pressure of 2×10⁻⁴ Torr=25 mPa.For example, a microwave discharge ion thruster was developed and testedat an operating pressure of 25 mPa. FIG. 14 shows that gas bulk velocityis about 20 m/s, reduced from the inlet velocity of 7.8 km/s and thatthe gas temperature is close to 400 K, which is manageable from a heatload perspective. By actively or passively cooling the duct walls, thegas temperature within the duct could be further controlled. Dependingon the characteristics of the C&S system, it is expected that the C&Ssystem will be capable of capturing up to about 40% of the incomingparticles and, in this example, channel them into the duct region. Givena nominal free-stream density of 1e-9 kg/m³ at 160 km altitude and 7.8km/s flight velocity with a cross-sectional area of 20 cm×45 cm for theC&S system this produces a captured mass-flow rate close to 0.3 mg/s.The calculated drag on the C&S system (opening or trap) using DSMC forthe 160 km example is approximately 5 mN, and for a C&S system and NEOvehicle combination (see FIGS. 1-5) is expected to be between 5-10 mN.Existing ion engines have been demonstrated to produce similar thrustlevels given similar gas conditions and flow rates.

The results discussed above show the potential for the C&S system toprovide a vehicle with thrust greater than drag. A first order argumentcan be made that, in this example, 60% of the particles bounce back outof the C&S system, but do so at the relatively low average velocity ofless than 300 m/s, compared to the incoming velocity of 7.8 km/s.Therefore, they impart drag nearly equal to the momentum they had whenthey struck the inner wall of the scoop section of the C&S system.However, the 40% of the particles concentrated and slowed by the C&Ssystem may be fed into an electric propulsion system that may acceleratethem to many keV of energy, or 25-400 times greater than the incomingenergy of the particles (typically arriving with about 10 eV and exitingwith anywhere from 250 eV up to a multiple thousand eV).

The described NEO vehicle 100 and incorporated C&S system provides forthe accelerated ions to impart substantially more thrust than thedecelerated (i.e., impacted) particles created in drag. In this case,assuming 1 kV of accelerating voltage, or about 100 times the kineticenergy of the impacted (drag producing) particles, only 6% of theincoming particles need to be accelerated to 1 kV to overcome the dragcreated by the 60% of the incoming particles that induced drag. This isbecause the 100 times increase in an exiting particle's energyrepresents a 10 times increase in momentum, or thrust.

The thermalized gas conditions created by the C&S system are suitablefor various forms of electronic propulsion devices (i.e., ion engines)such as pulsed plasma thruster (PPT), Hall-effect thruster (HET),microwave discharge, and RF discharge devices. An example microwavedischarge thruster was tested for an ionization chamber pressure of 25mPa (a pressure similar to that obtained by the C&S system) and produced30 mN of thrust using 1 kW of power. The thruster, however, operated onXe, an atomic species approximately four times more massive than theoxygen and nitrogen mix found in earth's upper atmosphere. Thrust in anion thruster is proportional to the square root of the mass of theaccelerated ion, so a skilled person will recognize that the C&S systemdescribed herein may provide conditions to generate about 15 mN ofthrust, still well above the drag created by the C&S system and examplevehicle geometry as described above.

An example ion engine was tested using air species (O₂, N₂, and a mix ofO₂ and N₂). The ionization chamber pressure of the engine is nominally2×10⁻⁴ Torr=25 mPa. The engine was tested for a range of conditions,however, the engine produced a nominal thrust of 5 mN using 450 W ofpower using air species. It is noted that this is close to the dragvalue for the example C&S system and vehicle geometry described withrespect to FIG. 1 and also may employ a solar panel area of just over 1m² to generate the 450 W of power, which is similar to the surface areaof the exemplary vehicle geometry (see FIGS. 1-5). This datademonstrates the feasibility of the SSIE with C&S system invention,potentially using existing ion engine technology. Similar concepts forlarger vehicles (cross-sectional area>1 m²), different deploymentconditions than those proposed herein, and customized ion enginedevelopment may allow for other successful implementations.

In other examples, the NEO vehicle 100 may incorporate navigation,cooling, attitude control, radio transmission, optical and radarimaging, power supplies, and digital processing. The resultingself-sustaining satellite can operate in long term, stable orbits ataltitudes between approximately 100-350 km, with the capability tocapture and transmit images of a given place on Earth on a highfrequency basis, be it hourly or even more often.

The endurance of the NEO vehicle 100 may exceed that of traditional ionengine powered satellites carrying their fuel into orbit. For example,Xe or Ar, both noble gases with relatively high atomic masses, can beused as a propellant. Noble gases are selected because they tend not todamage to elements of the engine, and massive atoms efficiently convertenergy into momentum. In the case of a NEO vehicle 100 with a C&S systemoperating with an ion engine, oxidation and sputtering by oxygen and tosome degree nitrogen may limit endurance of the system. Such issues maybe mitigated significantly, for example, by proper choice of materialsand by proper design of the ionization and acceleration chambers. Forexample, metallic elements, such as heavy, noble metals like gold do notoxidize and are less susceptible to sputtering than other materials. Newsynthetic materials or high strength ceramics may also be used.

The described NEO vehicle 100 that maintains a stable, self-sustainingorbit between 100-350 km can be part of an array of satellites in anorbital plane, defined as a satellite necklace (e.g., a single orbitalplane with multiple NEOs). Ninety NEOs in a single polar necklace willenable one of these satellites to traverse a given line of latitudeabout once per minute in a northbound direction assuming orbital timesof about 90 minutes, and again on the opposite side of the earth in asouthbound direction. In an example, twelve such satellite necklaces maybe arrayed, each separated by one hour of longitude, may be able toimage any spot on earth on average about once per hour.

In the example of FIG. 10, one or more NEO vehicles 100 can maintain anorbit 142 around the Earth 140, in accordance with the presentdisclosure. In one example, 90 satellites per necklace can be used,however more or fewer satellites per necklace may be appropriate for agiven application. For example, 45 satellites per necklace would spacethe vehicles at 2 minute intervals, while 180 would space vehicles at 30second intervals. As a person of ordinary skill will understand, theearth will rotate during the interval between arrival of two sequentialNEOs, with that distance determined by the time separation between thesatellites. Different spacing distances may impact other subsystemdesigns such as optical imaging and radio links, but the concept remainsthat a NEO satellite system can provide relatively high rates ofcoverage.

Since the time to revisit a same spot on earth is determined by the timefor that spot to rotate under the next necklace, doubling the number ofsatellite necklaces would reduce the revisit time for any spot on earthto 30 minutes or less, depending on the field of view of the onboardimager or radio. Conversely, halving the number of necklaces to 6 woulddouble revisit times. And reducing the number of necklaces to 4 wouldtriple revisit times. These changes would reduce system cost andcomplexity which may be a reasonable tradeoff for certain applications.

Other orbital planes can be utilized, and non-uniform distributions ofNEO vehicles could have beneficial applications. For example, non-polarorbits would increase the amount of time spent over populated areas andreduce the amount of total time spent over the poles. In the initiallydescribed system of 12 satellite necklaces with 90 NEO vehicles pernecklace, twelve NEO vehicles would fly over each pole every minute.Arrays of necklaces with inclinations to the equator of less than 90degrees could provide shorter revisit times for areas of greaterinterest.

Additionally, sequences of satellites with shorter distances betweenthem in a given necklace may be better suited for certain applications.For example, ten satellites separated by a few seconds could providesequential data on phenomena such as floods, fires or ice melting thatcould be useful in scientific understanding. A variety of NEO satellitedistributions are possible for various applications, each of which canemploy the NEO vehicle 100 described herein.

Different altitudes may be beneficial for certain applications. Widerangle coverage from higher altitudes may be an adjunct to higherresolution coverage from lower altitudes. Combination systems in whichSAR radar is combined with optical images may be desirable to operatethese different sensors (e.g. optical and radar) from differentaltitudes.

In another exemplary application, it may be beneficial to operate NEOsatellites operating at about 200 or 300 km to image orbiting space junkand satellites above them. Thousands of pieces of space junk, fromexpended launchers to small objects, represent a serious threat to LEOorbiting objects. Tracking such space junk from earth is difficult dueto their distance and atmospheric disturbance. Being much closer andmoving in independent orbits from those objects can improve trackingsubstantially.

For SAR radar applications, a smaller number of satellites per necklaceand a smaller number of necklaces may be sufficient to provide a desiredfrequency of useful images thanks to the all-weather and night imagingcapabilities of SAR radars. SAR radars employ multiple transmission andreception antenna arrayed in a specific pattern. The pattern istypically longer in one dimension (e.g. the direction of motion) than inthe transverse direction. Therefore, a relatively rectangular array ofelements may trail behind or be attached to a NEO vehicle to provide theoblong radiation beam needed to construct SAR images. By trailing orattaching such an array of antenna elements, drag will be impacted onlymarginally since it will be in the particle flow shadow of each NEOvehicle 100.

It is also possible to assemble arrays of NEOs positioned relative toeach other in a formation that may create the antenna array and beampattern needed for SAR. This may be an optimal approach for longerwavelength SAR applications since spacing between elements is typicallyrelated to the wavelength of the RF frequency being used.

The benefits and challenges associated with different orbits can beaddressed in response to a desired application's requirements.Combinations of SAR images with optical images provide uniquely usefulinformation. For example, radar may be able to image ground contoursthrough dense foliage that can be complemented by optical images of thefoliage. Different frequencies for both radar and optical imaging canalso add useful information. It is a benefit of the NEO system that nearsimultaneous imaging on a high revisit rate (e.g., hourly) providessubstantial improvements over traditional satellites at higheraltitudes.

In addition, a SAR-equipped satellite could be assigned to shadow anoptical imager, thereby providing tight correlation between radar andoptical images. Such a combination may provide a more comprehensiveunderstanding of activities on earth's surface than either type ofsatellite alone can offer. For example, a post-earthquake optical imagecan identify building damage that might be seen by terrestrial observerswhile radar images could highlight where vertical displacement hasoccurred in the building, or is occurring as a precursor to anaftershock. Such combinations today rely on long time lags between thetwo types of imagers, especially due to the scarcity of SAR-capablesatellites.

Orbital planes other than polar are obviously possible as well as ahybrid mix of polar and non-polar planes. The specific orbital plane maybe modified for different applications. Land mapping satellites may beconcentrated in lower latitudes since that is where the majority ofearth's land masses are found. The NEO vehicle 100 described herein canbe applied to any orbit in the targeted altitude, from polar toequatorial.

Short revisit times can be described above as “near-real time.”Traditional LEO and MEO satellites have revisit times of days to weeks,depending on the number of satellites in the constellation. Due toextremely high satellite costs plus high launch costs, satelliteconstellations are typically limited to a few to a few dozen satellites.Some proposed systems include up to about 100 satellites, promisingrevisit times down to a day or so.

Near real-time revisit rates offer many advantages and solve manyproblems inherent in current satellite systems. One example is the“worst case” revisit time as compared to the average revisit time. Mostsatellites spend about half their orbit in earth's shadow (i.e., night)resulting in poor or useless images. Adding in cloud cover, up to 70% ofearth's surface, sand storms and perspective issues (e.g., images takenaround noon cast no shadow and are therefore more difficult tointerpret) reduce the number of useful images to about one fifth or lessof all images taken.

This sampling problem makes it difficult to plan image capture of acertain spot at a certain time. For many implementations, the averagetime to a useable image may not be as important as the worst case time,which we define as the time between images that meet a certain set ofcharacteristics (e.g., a specific location plus morning or evening, plusno cloud cover, etc.). In this example, getting images of a specificarea (e.g., a battlefield or a river flood plain) with a long revisittime constellation can make a worst case scenario push from days intoweeks. In this example, a system with a 3-day average revisit time couldbe overhead at night for several sequential passes, and then encountercloud cover or dust storms when it is finally overhead with correctlighting. So an average revisit time of 3 days can become a one ortwo-week worst case scenario, a delay that reduces or even eliminatesthe value of the images.

Conversely, with an exemplary revisit time of an hour, the current NEOsystem will generally have a vehicle overhead any spot on earth duringdaylight hours, many times every day. Furthermore, as clouds and duststorms are not stationary, the probability of having a NEO vehicle 100overhead during a break in the weather is further increased. Since thesestatistics are not a purely linear extrapolation of the average revisittimes (i.e., they are exponential), worst-case revisit times become muchmore manageable with the described low revisit time NEO system.

Images are only useful once they are conveyed back to systems on Earth.The NEO vehicle 100 includes a widespread array of receiving stationsrather than the normally low number of centralized receiving stationsfound in use with traditional satellite systems. For example, with 3receiving stations (e.g., US, Australia and Europe), a traditional LEOsatellite will be within transmission range approximately every 30minutes (90/3), at best. If the imagery data is available with aninherent delay of a week due to the long revisit time described above, afurther 30 minute delay is relatively small.

However, for the current NEO satellite system, with average revisitrates of an hour or less, such a delay would be a large percentage ofthe goal. Therefore, data can be downloaded from the NEO vehicles to alarge network of low-cost earth receiving stations to enable low-latencydata downloads, ideally with latency from time of taking to time ofreceiving on the order of minutes to tens of minutes.

In one exemplary solution, receiving stations may be mounted atopcommercial cellular base stations, of which there are about 300,000 inthe US alone. Most such base stations are designed to support cellularcommunications radially outward. Therefore, an upwardly pointedradiation pattern can use the open area at the top of the base stationtower, directing and receiving all energy to/from an orbiting NEOsatellite and away from any interference with the cellular signals.

In order to download sufficient data during an overpass of a single NEOsatellite and to meet the size, mass and cost targets of the NEOsatellite, a simple antenna with a relatively wide beam will enable arelatively large footprint on earth's surface. For example, a beam withfull width half max (FWHM) beam angle of 45° from 100 km altitude wouldhave a circular footprint about 200 km in diameter. Assuming thevehicle's orbital velocity is about 7.8 km/s, a useable receive time ofabout 26 seconds would result. A narrower beam would reduce this timewhile a wider beam would increase it.

A tradeoff in the beam width is that as the beam width is reduced, themaximum data rate would typically increase. Hence a tradeoff is made tooptimize how much data can be downloaded during a single pass over agiven receiver. A further improvement can be made by using a more highlyfocused beam on the receiving site and having it track the NEO satelliteas it passes overhead. This would enable a relatively long dwell timedue to the relatively broad NEO satellite transmission beam, along withrelatively high data rates due to the relatively tight receiver beam.Also, since mass of the receiver is not as critical as mass on the NEOsatellite, placing a more complex (i.e., heavier) receiver and trackingantenna on the receiving side will reduce overall system cost.

In order to ensure low-latency downloads, downloads may occur when avehicle is passing over long stretches of ocean or other “dead zones”,of which the oceans are the largest. In addition to ensuringavailability of sufficient receiving stations on islands, receivers mayalso be placed on ships or buoys to receive the images which can then betransmitted to processing centers via traditional high capacitysatellites or fiber links.

In addition, NEO vehicles may include a vehicle-to-vehicle communicationsystem, such as with point-to-point laser systems. Since NEO vehiclesare designed to be as thin as possible to minimize drag, lasercommunications will be effective. In the example of 90 satellites in anorbital plane at 1 minute intervals, distance between satellites will beapproximately 450 km. Since the horizon from 160 km altitude is morethan 1,000 km away, a laser communications system is capable ofproviding a direct link to multiple satellites in the same orbitalplane. Since the vehicles will be oriented along the orbital plane inorder to minimize drag, the control system for the inter-vehicle lasercommunications may be simple, for example, including possibly a fixedorientation.

Using such an inter-vehicle link would enable very high-speed data ratetransfer between vehicles, enabling downloads to be handled by a vehicleother than the one collecting an image. Adding this flexibility to thesystem has several benefits, including filling dead-zone gaps, backupcapability if receivers are unavailable, and backup capability if adownlink transmitter on a NEO vehicle becomes disabled.

FIG. 15 provides a flow chart of an example method 150 of generatingthrust in a self-sustaining low earth orbit satellite, such as the NEOvehicle 100, in accordance with aspects of this disclosure. In step 152,ambient air particles are ingested, such as by the C&S system, to slowthe ambient air particles by at least two orders of magnitude and toconcentrate the ambient air particles by at least one order of magnitudewithin said NEO vehicle 100. In step 154, the ambient air particles areionized by an ion engine (e.g., ion engine 104). In step 156, theionized ambient air particles are accelerated within the ion engine 104of the NEO vehicle 100 and ejected to generate thrust.

In summary, a satellite system has been described that comprises thefollowing features:

-   -   Satellite operation in stable orbits at altitudes from 100-350        km;    -   Vehicle arrays sufficiently dense to enable overflights on        approximately hourly basis;    -   Concentrating & slowing system that concentrates and slows        incoming atoms and molecules to sufficient pressure and speed        enabling ion engine operation;    -   Solar powered, self-sustaining, air ingesting ion engines on        each vehicle providing thrust to counteract drag;    -   Thin, aerodynamic vehicles and payload shapes (i.e., thickness        less than width) that minimize drag at orbital velocities and        these altitudes;    -   High density array of receiving stations enabling low-latency        data downloads;    -   Laser-based vehicle-to-vehicle communications system;    -   Platform for optical and radar imagers; and    -   Self-cleaning orbital system for any generated space junk, to        name but a few advantages of the present disclosure.

A satellite system is described operating at altitudes between 100 and350 km relying on NEO vehicles that include a C&S System, feeding onatmospheric species into a SSIE to maintain near-constant orbitdynamics. The system operates at these altitudes that are substantiallylower than traditional satellites, thereby reducing the size, weight andcost of the vehicles and their constituent subsystems, such as opticalimagers and radio links. This reduction in size enables a virtuous cycleof further reduction in vehicle drag, which enables lower altitudeflight and further reduction in the size of vehicle components, etc.

The system includes a large number of the low-cost, low-mass, lowaltitude NEO vehicles, thereby enabling revisit times substantiallyfaster than any previous satellite system. The NEO vehicles spendvirtually all of their orbit at the low altitude, high atmosphericdensity conditions that have heretofore been virtually impossible toconsider. The C&S System conditions incoming atmospheric species to besuitable for use in ion engines, thereby enabling sufficient thrust toovercome drag. Short revisit times at low altitudes enable near-realtime imaging at high resolution and low cost. The system furtherincludes a distributed earth receiver system relying on a large numberof receivers each downloading data during a satellite overpass. Thecommunication link may utilize optimized beam shapes to maximize datadownload during each pass. A vehicle-to-vehicle laser communicationsystem may be included to improve data download rates, flexibility andreliability. By operating at such altitudes, the orbital mechanics haveno impact on the space junk issues of traditional LEO orbits and thesystem is self-cleaning in that any space junk or disabled craft willquickly de-orbit.

What is claimed is:
 1. A satellite network comprising a plurality of satellites arranged in an orbit having an altitude between 100 km and 350 km, wherein each satellite of the plurality of satellites comprises: a system configured to: ingest ambient air particles; transform the ambient air particles via a thermalizing, concentrating and slowing process; and an ion engine configured to: ingest the transformed air particles; ionize the transformed air particles; and generate thrust from the ionized transformed air particles sufficient to maintain the orbit of the satellite.
 2. The satellite network as defined in claim 1, wherein each satellite of the constellation of satellites is configured to at least one of: neutralize ionized transformed air particles downstream of the vehicle; and eject excess electrons.
 3. The satellite network as defined in claim 1, wherein each system comprises an inlet configured to: ingest the ambient air particles as the satellite moves in orbit; and thermalize, concentrate, and slow the ambient air particles.
 4. The satellite network as defined in claim 3, wherein one or more surfaces of the system comprise a material selected to be smooth relative to the size of the ambient air particles such that tailored focusing of the ambient air particles is achieved.
 5. The satellite network as defined in claim 4, wherein the material is sapphire.
 6. The satellite network as defined in claim 3, wherein one or more surfaces of the system comprise a material selected to transfer kinetic energy from the ambient air particles and to direct a portion of the ambient air particles into the inlet.
 7. The satellite network as defined in claim 6, wherein the material is selected based on particle scattering properties and is placed such that tailored focusing and thermalization of the ambient air particles is achieved.
 8. The satellite network as defined in claim 6, wherein the material is selected based on catalytic properties for converting O into O₂ and placed such that tailoring the amount of O and O₂ entering the ion engine is achieved.
 9. The satellite network as defined in claim 1, wherein the system is configured to decrease an average velocity of ambient air particles by at least two orders of magnitude and increase the pressure of the ambient air particles by at least one order of magnitude.
 10. The satellite network as defined in claim 1, each satellite further comprising a controller configured to: determine spatial information indicative of at least one of a current altitude of the satellite, an orientation of the satellite relative to a terrestrial surface, and a position of the satellite relative to at least one other satellite; compare a current altitude of the satellite against a desired altitude; and control the ion engine to generate thrust sufficient to achieve the desired altitude.
 11. The satellite network as defined in claim 1, each satellite further comprising a solar energy collection system to provide power to the ion engine and one or more components of the satellite.
 12. The satellite network as defined in claim 1, each satellite further comprising one or more sensors to determine the current altitude of the satellite.
 13. The satellite network as defined in claim 1, each satellite further comprising a radio transmitter and receiver to communicate with at least one of an airborne, a shipborne and a terrestrial based radio communication system.
 14. The satellite network as defined in claim 13, wherein each radio transmitter and receiver comprises a fixed antenna pattern.
 15. The satellite network as defined in claim 1, wherein the plurality of satellites occupy a first orbital plane, the satellite network further comprising a second plurality of satellites that occupy a second orbital plane different from the first orbital plane.
 16. The satellite network as defined in claim 15, wherein the first and second orbital planes comprise at least 45 satellites in each plane.
 17. The satellite network as defined in claim 15, wherein there are more than six orbital planes.
 18. The satellite network as defined in claim 1, each satellite further comprising a laser-based communication system configured to transmit information to and receive information from another laser-based communication system of another satellite.
 19. The satellite network as defined in claim 1, each satellite further comprising an active radar.
 20. The satellite network as defined in claim 1, each satellite configured to transmit information at near real time revisit rates relative to a terrestrial based receiver.
 21. The satellite network as defined in claim 18, wherein the plurality of satellites is arranged in an array sufficient to create a revisit time of an hour or less over a particular place on earth.
 22. A satellite configured to orbit a terrestrial surface at an altitude between 100 km and 350 km, the satellite comprising: a system with an inlet to thermalize, concentrate and slow incoming ambient air particles; and an ion engine configured to ingest and ionize the thermalized, concentrated, and slowed air particles and generate thrust from the ionized thermalized, concentrated, and slowed air particles sufficient to maintain the orbit of the satellite.
 23. The satellite as defined in claim 22, the system further comprising one or more thermalizing fins located within the inlet and arranged substantially parallel to the direction of travel, the one or more thermalizing fins configured to prevent the ingested ambient air particles from escaping in the upstream direction and to direct the ambient air particles into the inlet.
 24. The satellite as defined in claim 23, the channel configured to convey thermalized, concentrated, and slowed air particles to the ion engine, the ion engine to ionize the thermalized, concentrated, and slowed air particles and to accelerate and eject the ionized thermalized, concentrated, and slowed air particles from the satellite to generate thrust.
 25. The satellite as defined in claim 24, further comprising an active radar system configured to operate as a synthetic aperture radar.
 26. A method of generating thrust in a self-sustaining low earth orbit satellite, the method comprising: ingesting ambient air particles by a system configured to slow the ambient air particles by at least two orders of magnitude and to concentrate the ambient air particles by at least one order of magnitude within said satellite; ionizing the slowed and concentrated air particles by an ion engine; and accelerating the ionized slowed and concentrated air particles through an ejection port of said satellite to generate thrust.
 27. The method of claim 26, further comprising: collecting information from one or more sensors mounted on said satellite; and transmitting said information to a terrestrial receiver at a near real time revisit rate. 